Tsubame is a science and technology demonstration mission developed by a student-lead system design of the Matunaga LSS (Laboratory for Space System) at the Tokyo Institute of Technology (referred to as TITech or simply Tokyo Tech). Tsubame is the fourth satellite project of TITECH, after CUTE-1 (launch June 30, 2003), Cute-1.7+APD (launch Feb. 21, 2006), and Cute-1.7+APD II (launch April 28, 2008). 1)2)

The name Tsubame means "swift" in Japanese, referring to the fast spacecraft attitude changes for event monitoring; it also connotes a certain relationship to NASA's gamma-ray mission, Swift. In addition, Tsubame happens to be the logo of Tokyo Tech, symbolized by a swallow.

• A goal is to demonstrate the satellite bus technology for the 50 kg microsatellite class and to verify COTS (Commercial Off The Shelf) components such as microprocessors, memory, and Li-ion batteries in the space.

• The second mission goal is the high-speed attitude maneuver experiment using newly developed CMGs (Control Moment Gyros) to demonstrate spacecraft agility.

• Demonstration of a high-resolution (14 m) optical camera as part of the payload.

2) Science and engineering mission - involving the observation of polarized X-ray and gamma-ray radiation sources. The nature of GRBs (Gamma Ray Bursts) is generally of short duration (events on the order of 1 minute); in addition, GRBs happen to be unpredictable events to occur anywhere within the celestial sky. Hence, an agile platform is required to permit proper observations after detection of a GRB source. The GRB instrument is called WBM (Wide-field Burst Monitor).

The Tsubame mission is a cooperative project involving the following partners:

- The LSS team at the Tokyo Institute of Technology (TITECH) is responsible for the management of the project, development, test, launch, operation as well as for the development of the satellite bus system.

- The gamma-ray observation sensors are developed by the Tokyo Institute of Technology, Kawai laboratory.

- The Tokyo University of Science Kimura laboratory is responsible for the development of the optical camera module.

- In addition to these laboratories, LSS collaborates with two companies: Tamagawa Seiki Co., Ltd. develops the CMG hardware. SpaceLink Co. Ltd. helps the Tsubame project in terms of the development of EPS (Electrical Power Subsystem) and the OBC (Onboard Computer).

As of fall 2010, the development of the Tsubame microsatellite is in the EM (Engineering Model) development phase. The Tsubame project is sponsored by MEXT (Ministry of Education, Culture, Sports, Science and Technology) of Japan in the context of "small satellite research and development projects."

Tsubame is a microsatellite with an almost cubic platform of size 45 cm x 45 cm x 56 cm (H) and a total mass of ~ 49 kg. The S/C bus uses a lattice structure (CFRP material). Each of the four bus panels is referred to as Bulkhead. The structural design of the bus has been subjected to vibration tests at JAXA. The design life of the satellite is 1 year. The spacecraft is 3-axis stabilized (Ref. 1). 6)7)8)

Thermal and structural subsystems: The spacecraft structure offers such functions as a paddle deployment and separation mechanisms. The passive thermal subsystem provides the required environment for all subsystems.

EPS (Electrical Power Subsystem): Use of four deployable SAPs (Solar Array Panels) and some solar cells (InGaAs) which are mounted on the bus surfaces. The SAPs are needed in support of the high power (130 W max, 86 W average) requirements of the mission. The power bus voltage is 32 V when in sun-pointing mode, and 25.2 V (max) in general support modes. A Li-Po (Lithium-Polymer) battery (6 cells) is used in the eclipse phases of the orbit. Use of sequential shunt control. The EPS along with the battery protection circuit as well as the OBC are controlled by the PIC microcontroller (Figure 8).

C&DH (Command and Data Handling) subsystem: The OBC (On-Board Computer), a Power PC on FPGA, in charge of all C&DH functions including mission sequence control and data processing. A PIC (Peripheral Interface Controller) processor is in charge of monitoring the OBC.

The payloads and each subsystem, except the EPS, feature an FPGA-based microcontroller; they are interfaced with the CAN (Controller Area Network) bus. For redundancy, each subsystem is connected through a UART (Universal Asynchronous Receiver/Transmitter) link with Comm/C&DH subsystems - in case any communication failures on the CAN bus.

Figure 9: Layout of the C&DH subsystem (image credit: TITech)

RF communications (also referred to as Comm): The amateur bands (UHF, VHF) are used for TT&C data transmissions (AX.25 protocol). The UHF (430 MHz) band is for the downlink and VHF (144 MHz) band for the uplink. In addition, a CW beacon is used (GMSK 9.6 kbit/s / AFSK 1.2 kbit/s). The payload data are transmitted in S-band at a max. data rate of 100 kbit/s (BPSK modulation).

Launch: The Tsubame microsatellite was launched as a secondary payload on November 6, 2014 (07:35:49 UTC) on a Dnepr-1 vehicle from the Yasny Cosmodrome, Russia. The primary payload on this flight was the ASNARO minisatellite of USEF, Japan.
The launch provider was ISC Kosmotras. The launch was executed by the Russian Strategic Rocket Forces of the Russian Ministry of Defense with the support of the Russian, Ukrainian and Kazakhstan organizations, which are part of the ISC Kosmotras industrial team. 14)

1) Detumbling mode (occurs after spacecraft deployment). It is also being used whenever the spacecraft should become unstable or when an FDIR (Failure Detection, Isolation and Recovery) problem arises in any subsystem.

2) Safe mode: This is a sun-pointing mode with the objective to load the spacecraft batteries.

3) Standby mode: A nominal mode used prior to an observation engagement (also used for control of the spacecraft). The standby mode is used to check out the spacecraft subsystems and to test for instance the capabilities of the new CMG device.

4) Camera mode: Used in support of landmark tracking to collect imagery of particular target regions.

5) Camera mode: Used in support of panorama observations.

6) Gamma-ray mode: Used in support of GRB (Gamma-Ray Burst) observations.

Mission status:

• April 2015: TSUBAME stopped receiving uplink commands during the initial checkout phase. Location of the faults : the peripheral circuit for communication is suspicious. 16)

1) The communications rate can be more than 10 kbit/s usually, though it is difficult to achieve 100 kbit/s stably

2) TSUBAME was able to endure the Dnepr rocket vibrations

3) TSUBAME was able to point to the sun and balance electrical power within 33873 seconds after rocket separation, and the maximum depth of discharge was estimated at no more than 50%. In the flight operation, TSUBAME automatically deployed the solar array paddles after the rocket separation; then, TSUBAME performed the sun pointing autonomously and supplied electricity stably for at least 80 days; moreover, the temperature profile of TSUBAME agrees well with the thermal analysis. The failure analyses and design improvements are in progress.

- On Nov. 7, 2014, the 430 MHz-band Tx1, the 430 MHz-band Tx2, an S-band Rx and a GPS receiver were turned on in a checkout phase. However, the condition of TSUBAME got worse after November 12. An S meter indicated an abnormal value, that is, an 430 MHz-band receive function failed. Then, TSUBAME entered the safe-hold mode on December 12, while S-band uplinks were prepared in the ground station. Hence, TSUBAME lost receive function because the S-band receiver was powered off. Finally, the CW was no longer heard from since January 27, 2015. This caused a loss of contact with the ground station. Failure analyses and design improvements are in progress.

- TSUBAME had to deal with depletions of its batteries after the rocket separation. This was due to the power shortage of solar arrays while the sun angle (error angle between sun direction and a vertical direction to solar array paddles) is over 67º . Therefore, TSUBAME had to point to the sun as soon as possible after the separation (Ref. 17).

- A HiLS (Hardware-in-the-Loop Simulation) testing with an attitude simulator and a solar array simulator was conducted in order to predict a battery voltage, or a depth of discharge, during the initial operation. The longest estimated time to point to the sun was 33873 seconds after the separation. The experimental setup is shown in Figure 13.

- It was ensured that TSUBAME could survive the power crisis leaving at least 50% margin of the depth of discharge. The lowest battery voltage during the experiment was 23.2 V as shown in Figure 4. The battery voltage means the depth of discharge was 50%. In fact, the depth of discharge was no more than 50% in the flight operation.

• On the first pass over Japan, the TSUBAME development team successfully received telemetry data from TSUBAME at the TITech (Tokyo Institute of Technology) ground operation station which confirmed that the satellite was operating normally. It was also confirmed that all the solar array panels were deployed, and that it is operating under spin stabilized attitude control pointing towards the sun. 18)

- Approximately 15 minutes after launch, TSUBAME was separated from the rocket in airspace above the Indian Ocean.

Experiment/Sensor complement: (CMG, Optical Camera, GROS)

CMG (Control Moment Gyroscope)

The CMG is a joint development project of LSS (software) and Tamagawa Seiki Co. (hardware). The design of the CMG assembly consists of four CMGs configured in a pyramid fashion. 19)20)21)

Note: A CMG is an attitude control actuator that gives the necessary torque to the satellite by exchanging the angular momentum between the rotating wheel and the satellite. A CMG can produce larger torque than conventional actuators such as reaction wheels. By combining a characteristic of a small satellite, which has a relatively small mass, and high torque that the CMG device can generate, agile maneuvering can be realized.

The small optical camera module is developed by the Kimura laboratory at the Tokyo University of Science. Tsubame features a high-resolution optical camera, which has a size of 90 mm x 100 mm x 200 mm, a mass of 1kg and a spatial resolution of 14 m from an orbit of 500 km altitude. Use of COTS components to keep the development within budget (Ref. 1).

The goal of the optical camera is to provide the following EO (Earth Observation) missions: One is landmark tracking to collect sufficient information of target regions of interest. The other one is panorama shooting service to obtain broad imagery from the high resolution camera, which has a narrow FOV (Field of View).

The microprocessor of the camera system is based on an FPGA (Field Programmable Grid Array) implementation; this concept provides flexibility to adapt easily to different types of satellite bus systems. On the software front, Linux is adopted for the camera system to permit an adaptive environment for functional enhancements.

Instrument size, power, mass

90 mm x 100 mm x 200 mm, < 3 W, 1 kg

Detector array

2,210 x 3,002 pixels

Pixel size

3.5 µm x 3.5 µm

Focal length

175 mm

Full frame rate

5 frames/s

Table 3: Specification of the optical camera

GROS (Gamma-Ray Observation System):

The objective is to provide GRB (Gamma-Ray Burst) observations with the agile Tsubame spacecraft. The X-ray detector system consists of two detectors: WBM (Wide-field Burst Monitor) and HXCP(Hard X-ray Compton Polarimeter). Both detectors have been developed by the Kawai laboratory at the Tokyo Institute of Technology (Ref. 1). 22)23)24)

The system configuration of gamma-ray observation equipment is shown in Figure 20.

The WBM device functions as the GBR detector while the HXCP device functions as the observer. Both devices (WBM and HXCP) feature a scintillator APD (Avalanche Photodiode) design.

The objective of WBM is to detect a burst occurrence and to determine the event (GRB transient) direction within a beam width of ±10º. When a GRB is detected, the onboard CPU calculates the coordinate and then the satellite slews rapidly using the CMG (Control Moment Gyro). Thanks to the high speed attitude control system, Tsubame can start polarimetry observations earlier than ~15 s from the trigger.

The WBM is always monitoring half the sky and observing the count rate of hard X-ray photons. When a GRB is detected as rapid increase of the count rate, the onboard CPU calculates the coordinate of GRB by comparing count rates among five counters and then the satellite slews rapidly (Figure 21).

The WBM employs five gamma-ray counters that consist of the CsI (Tl) scintillator (size: 12 cm x cm, thickness: 5 mm) and APD. Each counter is mounted on the surface of the satellite. These detectors monitor the count rate of the X-ray and gamma-ray photons and trigger if the count rate rises rapidly. In order to determine the position of the transient object, the detectors are pointed into five different directions. By comparing the count rate of each detector, the source position can be estimated by using the barycentric method. This kind of technique had already been applied to the full-scale gamma-ray observing missions, such as the BATSE (Burst and Transient Source Experiment) aboard the CGRO (Compton Gamma-Ray Observatory) or the GBM (Gamma-ray Burst Monitor) aboard the Fermi observatory.

Detectable count rates

> 10 Hz

Detectable minimum energy

< 30 keV (at 20ºC)

Accuracy of position determination

< 5º

Device mass

260 g / unit

Table 4: Performance goal of WBM

HXCP (Hard X-ray Compton Polarimeter)

The HXCP requirements call for the polarimeter to observe the GRB event (i.e., point into the proper direction) within 10-20 seconds after the event detection by the coarse calibration sensor. The goal of HXCP is to detect photons in the 30 – 200 keV band using the azimuthal angle anisotropy of Compton scattering. 25)

The HXCP device consists of scatterer for scattering the X-rays and determining the incident position, and absorber for detecting the scattered photons and measuring the scattering angle and the energy of the photons. The scatterer consists of 64 plastic scintillators (6.5 mm x 6.5 mm x 49 mm) placed at the center of the detector. The absorber is comprised of 28 CsI (Cesium Iodide) scintillators (6.5 mm x 10 mm x 49 mm). Their signals are read by the MAPMTs (Multi-Anode Photo Multiplier Tubes) and the Si APDs (Avalanche Photodiodes), respectively. A special 16 channel MAPMT was developed for the Tsubame mission which provides quite a high quantum efficiency; it uses mechanically strengthened electrodes that survived the H-IIA QT level.

The APD design employs the the recently progressed optical radiation detectors, which possess excellent properties: compact, lightweight, mechanically hard, low power consumption, and high quantum efficiency, resembling the other kind of semiconductor detectors. In addition, the unique property of the APD is the internal signal multiplication caused by the electron cascades. Thanks to this property, the APD has a good signal to noise ratio at room temperature.

For the detection of X-ray polarizations, it is essential to reduce the background events caused by the cosmic X-ray backgrounds, atmospheric gamma rays, and by trapped charged particles. In order to reject those events, a coincidence technique is applied. Since an expected Compton event will hit the scatterer and absorber at the same time, the signal processor works only when the MAPMTs and the APDs detect a signal simultaneously.

By selecting the hit patterns, the device can distinguish whether the event is a photon from GRB or an accidental noise. Using high performance and ultra low-power analog VLSIs that can processes multi signals simultaneously, the signal processing system could be downsized. The geometrical structure of the HXCP was designed based on Monte-Carlo simulations to achieve high sensitivity in spite of its small effective area.

Legend to Figure 22: The scintillators and detectors are accommodated in an upward direction, the passive shields surround them. The circuit boards and power supply units are placed under the detectors.

The information compiled and edited in this article was provided byHerbert J. Kramer from his documentation of: "Observation of the Earth and Its Environment: Survey of Missions and Sensors" (Springer Verlag) as well as many other sources after the publication of the 4th edition in 2002. - Comments and corrections to this article are always welcome for further updates (herb.kramer@gmx.net).